One of the most critical problems in small gas turbines, such as for missiles, trainers, small military or business aircraft, general surface transportation and ground power, is the difficulty associated with achieving strong and efficient turbine blade cooling with reliable, simple and inexpensive turbine blade structures. In such gas turbines, the turbine blades are usually too small to employ the turbine blade cooling techniques as used for large gas turbines.
The growing need for effective blade cooling techniques in small gas turbines is a direct consequence of the continuous quest for greater fuel economy. Greater fuel economy necessitates higher compressor pressure ratios and correspondingly higher isentropic compressor exit temperature (T.sub.c,e,s). In turn, the turbine inlet temperature (TIT) must be raised, since the thermal efficiency of the Brayton cycle, with non-ideal components, begins to drop rapidly when the ratio of TIT/T.sub.c,e,s decreases below 2. For example, for a compressor pressure ratio of about 30:1, T.sub.c,e,s /T.sub.amb =2.64: thus TIT/T.sub.amb =5.28, and TIT=2750.degree. R. at an ambient temperature (T.sub.amb) of 520.degree. R.
Current attempts to make ultrahigh, turbine-inlet gas temperatures possible are based on the use of turbine materials having ultrahigh temperature and stress capabilities such as advanced ceramics or advanced composite materials, or the use of oxidization resistant coatings for exotic turbine-blade materials such as Columbium and others. Solving the high-temperature problems by the use of uncooled, novel materials entails many risks and uncertainties with respect to development time, cost and ultimate reliability.